Rolling missile guidance system having body fixed antennas



Aug. 11, 1970 3,523,659

ROLLING MISSILEGUIDANCE SYSTEM HAVING BODY FIXED ANTENNAS E. H.l EPPERSON, JRv

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ROLLING MISSILE GUIDANCE SYSTEM HAVING BODY FIXED ANTENN'AS ,HDW/Af Mw/esa@ #fram/,5%

Aug. 1l, 1970 y E. H. EPPERSON, JR

ROLLING MISSILE GUIDANCE SYSTEM HAVING BODY FIXED ANTENNAS Filed March 4, 1968 4 Sheets-Sheet 4 w w @u i Nmmfwlv a,

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United States Patent O 3,523,659 ROLLING MISSILE GUIDANCE SYSTEM HAVING BODY FIXED ANTENNAS Edwin H. Epperson, Jr., Upland, Calif., assignor to General Dynamics Corporation, a corporation of Delaware Filed Mar. 4, 1968, Ser. No. 710,078 Int. Cl. F41g 9/00; F24b 15/02; H0111 1/28 U.S. Cl. 244-3.17 2 Claims ABSTRACT OF THE DISCLOSURE In a rolling missile, a guidance system which utilizes a pair of antennas in a single plane plus associated circuitry to direct the rolling missile towards a preselected target. g

BACKGROUND OF THE INVENTION In a guided missile, various receptors can be utilized to receive reected electromagnetic energy from a target. The position of the intended target relative to the guided missile can be measured in terms of any conventional three-dimensional coordinate system. Normally, however, spherical coordinates are used, that is, range, azimuth, and elevation. Of these the azimuth and elevation angles are primarily used to determine the direction in which the missile ight path should proceed.

Traditionally these two coordinates are measured either by scanning a single antenna such that its main lobe will traverse a circle around some directed line in space passing through the missile axes but not necessarily through the target or, alternately, they are measured by some form of monopulse system which measures both azimuth and elevation either simultaneously or alternately through either one time shared channel or two separate channels, In either case, the antennas must be capable of sensing the direction of the source of the reected energy in at least two planes relative to the missile body. These traditional systems have many inherent disadvantages.

SUMMARY OF 'IHE INVENTION The invention resolves the problems of the prior art by providing a pair of single plane body fixed antennas in a rolling missile. The signals received from the antennas are phase shifted to provide -an error signal from associated circuitry to direct the missile towards its intended target.

Therefore, it is an object of this invention to provide a rolling missile having a pair of single plane body fixed antennas.

Another object of the invention is to provide a pair of single plane body fixed antennas for a rolling missile wherein a single information channel is employed between the antennas and the missile control surfaces.

Another object of fthe invention is to provide a pair of single plane body fixed antennas for a rollin-g missile in which the missile airframe, the antenna angle sensing plane, and the missile control surfaces are rolled in synchronization.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a geometric representation of the rolling missile with respect to an associated target.

FIG. 2 is a single plane vector representation of the rolling missile with respect to an associated target.

lFIG. 3 is a front-end view of the rolling missile and its body fixed antenna.

FIG. 4 is a top view of the rolling missile of FIG. 3.

PIG. 5 is a side view of the rolling missile of FIG. 3.

3,523,659 Patented Aug. 1l, 1970 FIG. 6 is a schematic diagram of the body xed antenna and associated circuitry.

DESCRIPTION `OF THE PREFERRED EMBODIMENTS The guidance system of the present invention is designed for utilization in a rolling missile such as that described in U.S. Pat. No. 3,333,790 and 3,351,303. The operation and control of such missiles is described in considerable detail in the above two patents and is wellknown in the art.

Referring now to IFIG. 1 there is shown a rolling missile airframe '10 rolling about an axis 15 and having a pair of antennas 12 and 14. The antennas I12, and 14 are mounted at diametrically opposed positions at the front end of the rolling airframe, and the plane of the two antennas (antenna angle sensing plane) is represented as plane 16. Line 18 represents the electrical boresight of the antennas and is displaced at angle cosine qs from the missile roll axis 15. A target 22 in the missile Ioll/ target plane 24 is located at an angle ,B plus e from the rolling airframe roll axis d5. 'Ihe missile to target line of sight 26 extends between the rolling airframe 10 and the target 22 and is projected perpendicularly onto plane 16 as line 20 located an angle e cosine q from the electrical boresight 18.

The same target 22 and missile y10 having antennas 12 and 14 is shown in FIG. 2, but in a single plane vector ge'ometry. Target 22 moves with a target velocity vector 28 in the direction indicated by the arrow. The missile position vector 30 (antenna mechanical boresight) is displaced a counterclockwise angle 1p from a space reference 32. The antenna electrical boresight 34 is displaced at angle from the missile position vector 30 or an angle plus 5b equals 0 from the space reference. The missile to target line of sight 36 is displaced a counterclockwise angle e from the antenna electrical boresight 34 or angle @-f-e equals ofrom the space reference 32. The missile velocity vector 38 is displaced a counterclockwise angle y from the space reference 32.

In FIG. 3 the front end of the missile or rolling airframe 10 is shown having two surface waves, circularly or linearly polarized dielectric-rod antennas 12 and 14 diametrically opposed at the front end of the missile 10. The rods may be made from polystyrene or other good dielectric material. The missile is shown rolling in a coun terlockwise direction as indicated by -the arrow but may likewise be rolled in a clockwise direction. The target 22 is situated at an angle '90--q5 from the angle sensing plane of the antennas -12 and 14.

FIG. 4, an instantaneous top Iview of the missile 10 FIG. 3 exposes only ythe top antenna 12 whereas FIG. 5, an instantaneous side View of FIG. `3 illustrates both antennas |12 and -14 with the target located an angle (-i-e) sine qs from the axis of rotation of the missile.

The schematic diagram of FIG. 6 includes both antenna 12 and antenna 14. Two RF mixers 40 and 42 coupled to a common phase shifter 44vfed by a local oscillator 46 receive the output from antennas 12 and 14 respectively. The outputs from RF mixers 40 and 42 are both fed to a phase detector 48 which is connected back to the phase shifter 44 through an amplifier 50 and an integrator 52.

In operation, the antennas 12 and 14 receive the same energy from the target 22 and therefore produce the same information only differing by the period of time and thus phase, depending on the angle of incidence of the line of sight from the missile 10 to the target 22 relative to the mechanical boresight of the two antennas 0 or missile roll axis. This angle of incidence varies sinusoidally as the missile rolls and its maximum variation is ,S-i-f, thus the electric phase difference a between the two antennas relative to the incoming energy is described by:

azZnD, sin l(+e) cos (tmf-10] where:

(,B-I-E)-the maximum angle of incidence of the received energy with respect to the missile roll axis,

wR-roll rate of the missile airframe in radians per second,

lb-polar angular displacement of target relative to a reference plane through the missile airframe roll axis (FlG. l), and

D-separatiom in wavelengths of received energy, be-

tween the two antenna elements.

Using a small angle approximation for sine (t3-H),

rx=21rD,\(-le) COS (Rf-q) The phase, alg or am, of the signal at each antenna, with respect to a point midway between the two antennas, is either as illustrated in the phase tracking loop of FIG. 6.

Either one or both of the RF channels may contain phase-Shifters controlled by a common feedback loop. For purposes of illustration a single phase shifter 44 is shown in FIG. 6. The action of the feedback loop is such that the phase differences between the outputs of the two channels tends to zero. lf the missile to target line of sight changes for example, one degree, the feedback loop attempts to shift the phase of one or both channels to bring the net phase difference of the two channels back to zero. In FiG. 6 the phase-shifter 44 is shown driving two summing points 40 and 42 equally in amplitude and opposed in phase. The amplitude of these feedback signals denoted by a'lz and 1'14 will be proportional to and in fact they define ,3, since the electrical boresight ,B is determined by the relation between @U12 and '14. If a'lgza'm then {3:0, The summing points in this case consist of a pair of RF mixers for heterodyning the input frequencies down to convenient intermediate frequencies and the phase shifter(s) are placed in series with lines from a common local oscillator.

In addition, a relatively high frequency sinusoidal phase modulation commonly referred t as scanning may be superimposed on each loop of the phase shifter (equally in magnitude and opposed in phase) after which the two cycles can be summed. The particular signal processing scheme is however, not important in that there are many alternate circuit arrangements available to accomplish the same purpose.

The output of the summing devices 40 and 42 consists of an intermediate frequency as a carrier with an amplitude modulation containing scan-plus-roll and scanminus-roll frequency sidebands. As the missile rolls the value of E in turn varies sinusoidally at a roll frequency equal to the missile roll rate and therefore the magnitude of the scan modulation varies sinusoidally at a rate equal to the missile roll rate resulting in the sidebands described above. The scan frequency may be removed later in a demodulator and the derived signal will be principally a sinusoidal voltage (with respect to time) whose frequency is equivalent to the missile roll rate, whose amplitude is proportional to the net phase difference between the two RF channels, and whose phase is proportional to the angle determined by the direction of the line of sight error causing the phase difference. The phase detector 4S shown schematically would include the summation device, IF amplifiers, video circuitry, the scan demodulator and appropriate lters. The output of the phase detector is then amplified an appropriate amount.

The integrator 52 may be a synchronous lter type i (which demodulates to DC., integrates with an operational amplilier, then modulates back to roll frequency) keyed by a rollo accelerometer. The output of the integrator 52 then drives the phase shifter 44 sinusoidally (at the missile roll rate).

The error signal or output amplitude is obtained from a point between the amhplier 5t) and integrator S2. It can however be taken at a point further back in the feedbackdoop. This output Ke cosine (WRt-) is equivalent to a cosine (WRI-tp) since in a closed loop such as this having a single integrator, the amplitude Ks of the output is a measure of the rate of change of the mis- Sile to target line of sight. This output can then be used to drive the missile control system and thereby provide a sinusoidal (at the missile roll rate) variation in the missile velocity vector fy.

As can be seen, a single information channel can be utilized the entire distance between the antennas and the missile control surface. Since the missile airframe, the antenna angle sensing plane, and the control surfaces are rolling in synchronization, this results in a multifold reduction in simplification of hardware and electronic signal processing when compared to conventional antenna and signal processing receiver and guidance systems.

The sinusoidal error signal will be in phase synchronization with the hypothetical plane paced through the missile roll axis and rigidly attached to the airframe and its associated control surfaces. Since this error signal is sinusoidal and contains both the phase and magnitude information descriptive of the exact direction and magnitude of rate of change of the missile to line of sight, the control signal of each control surface varies sinusoidally at the missile roll rate thus forcing the mean missile velocity rate of change in a direction parallel to the missile to target line of sight rate of change. The result is proportional navigation, by definition, and an eventual intercept of the target by the missile.

What is claimed is:

1. A rolling missile guidance system comprising:

(a) first receptor means fixedly mounted upon the :rolling missile to receive energy from a prospective target and produce a first output signal;

(b) second receptor means fixedly mounted upon the rolling missile in the same transverse plane as said first receptor means to receive energy from the same prospective target and produce a second output signal differing in phase from said first output signal of said first receptor means; and

(c) signal processing means opel-ably associated with said first and said second receptor means to receive said rst and said second output signals therefrom to produce a proportional navigational error signal to direct the rolling missile towards an intercept with the prospective target;

said signal processing means comprising:

(d) first summing point means operably connected to said first receptor means to receive said first output signal from said first receptor means;

(e) second summing point means operably connected to said second receptor means to receive said second output signal from said second receptor means;

(f) phase shifter means operably connected to said first and second summing point means for driving said first and second summing point means equally in amplitude and opposed in phase;

(g) an oscillator operably coupled to said phase shifter means;

(h) phase detector means operably connected to said first and second summing point to receive output signals from said first and second summing points;

(i) an amplifier operably connected to said phase detector means and responsive thereto; and

(j) integrator means operably connected to said amplifier for producing output signals that are the integrals of input signals from said amplifier.

6 2. The rolling missile guidance system of claim 1 Where- 3,363,858 1/ 1968 Dobbins et al. 244-3.14 in said rst and said second summing point means are 3,405,888 10/ 1968 Okamoto 244-3.15 RF mixers.

References Cited BENJAMIN A. BORCHELT, Primary Examiner UNITED STATES PATENTS 5 T. H. WEBB, Assistant Examiner 2,867,776 1/ 1959 Wilkinson 343-911 2,929,065 3/ 1960 Kreinheder 343-911 U'S' C1' X'R' 3,001,186 9/1961 Baltzer 244-315 343-705 708 911 

